Geared turbofan engine with targeted modular efficiency

ABSTRACT

A turbofan engine includes a fan section including a fan blade having a leading edge and hub to tip ratio of less than about 0.34 and greater than about 0.020 measured at the leading edge and a speed change mechanism with gear ratio greater than about 2.6 to 1. A first compression section includes a last blade trailing edge radial tip length that is greater than about 67% of the radial tip length of a leading edge of a first stage of the first compression section. A second compression section includes a last blade trailing edge radial tip length that is greater than about 57% of a radial tip length of a leading edge of a first stage of the first compression section.

CROSS-REFERENCE TO RELATED APPLICATION

This application is a continuation of U.S. patent application Ser. No.16/531,704 filed on Aug. 5, 2019, which is a continuation of U.S. patentapplication Ser. No. 14/651,923 filed on Jun. 12, 2015, now U.S. Pat.No. 10,371,047 granted on Aug. 6, 2019, which is a National PhaseApplication of International Application No. PCT/US2014/057127 filed onSep. 24, 2014, which claims priority to U.S. Provisional Application No.61/891,475 filed on Oct. 16, 2013.

BACKGROUND

A gas turbine engine typically includes a fan section, a compressorsection, a combustor section and a turbine section. Air entering thecompressor section is compressed and delivered into the combustionsection where it is mixed with fuel and ignited to generate a high-speedexhaust gas flow. The high-speed exhaust gas flow expands through theturbine section to drive the compressor and the fan section. Thecompressor section typically includes low and high pressure compressors,and the turbine section includes low and high pressure turbines.

The high pressure turbine drives the high pressure compressor through anouter shaft to form a high spool, and the low pressure turbine drivesthe low pressure compressor through an inner shaft to form a low spool.The fan section may also be driven by the low inner shaft. A directdrive gas turbine engine includes a fan section driven by the low spoolsuch that the low pressure compressor, low pressure turbine and fansection rotate at a common speed in a common direction.

A speed reduction device such as an epicyclical gear assembly may beutilized to drive the fan section such that the fan section may rotateat a speed different than the turbine section so as to increase theoverall propulsive efficiency of the engine by allowing an increase inthe fan diameter and a reduction in a fan pressure rise. In such enginearchitectures, a shaft driven by one of the turbine sections provides aninput to the epicyclical gear assembly that drives the fan section at areduced speed such that both the turbine section and the fan section canrotate at closer to their individual optimal speeds.

Although geared architectures have improved propulsive efficiency,turbine engine manufacturers continue to seek further improvements toengine performance including improvements to thermal, transfer andpropulsive efficiencies.

SUMMARY

A turbofan engine according to an exemplary embodiment of thisdisclosure, among other possible things includes a fan section includinga fan blade having a leading edge and hub to tip ratio of less thanabout 0.34 and greater than about 0.020 measured at the leading edge anda speed change mechanism with gear ratio greater than about 2.6 to 1. Afirst compression section includes a last blade trailing edge radial tiplength that is greater than about 67% of the radial tip length of aleading edge of a first stage of the first compression section. A secondcompression section includes a last blade trailing edge radial tiplength that is greater than about 57% of a radial tip length of aleading edge of a first stage of the first compression section.

In a further embodiment of the foregoing turbofan engine, the fansection provides a low fan pressure ratio less than about 1.6.

In a further embodiment of any of the foregoing turbofan engines, thefan section provides a low fan pressure ratio between about 1.45 andabout 1.20.

In a further embodiment of any of the foregoing turbofan engines, thefan section provides a bypass ratio greater than about 8.

In a further embodiment of any of the foregoing turbofan engines, thefan section provides a bypass ratio greater than about 8.

In a further embodiment of any of the foregoing turbofan engines, thefan section provides a bypass ratio greater than about 12.

In a further embodiment of any of the foregoing turbofan engines,includes a turbine section that has a fan drive turbine and at least twoturbine stages forward of a first turbine blade of the fan driveturbine.

In a further embodiment of any of the foregoing turbofan engines, thefan drive turbine includes at least three stages.

In a further embodiment of any of the foregoing turbofan engines, thefan drive turbine is coupled to the drive the first compression section.

In a further embodiment of any of the foregoing turbofan engines, atleast one of the at least two turbine stages is coupled to drive thesecond compression section.

In a further embodiment of any of the foregoing turbofan engines, the atleast two stages include a single turbine second forward of the fandrive turbine.

In a further embodiment of any of the foregoing turbofan engines, thefirst compression section includes three stages and the secondcompression section includes eight stages.

A turbofan engine according to an exemplary embodiment of thisdisclosure, among other possible things includes a fan section providinga bypass ratio greater than about 12, and a speed change mechanism withgear ratio greater than about 2.6 to 1. A first compression sectionincludes a last blade trailing edge tip length that is greater than 67%of the radial length of a first stage leading edge of the firstcompression section. A second compression section includes a last bladewith a trailing edge tip that includes a radial length that is greaterthan 57% of a radial length of the first stage leading edge of the firstcompression section.

In a further embodiment of any of the foregoing turbofan engines, thefan section includes a plurality of fan blades supported on a hub. Aleading edge of at least one of the fan blades includes a leading edgeand a hub to tip ratio is less than about 0.34 and greater than about0.020 measured at the leading edge.

In a further embodiment of any of the foregoing turbofan engines,includes a turbine section that has a fan drive turbine and at least twoturbine stages forward of a first turbine blade of the fan driveturbine.

In a further embodiment of any of the foregoing turbofan engines, atleast one of the at least two turbine stages is coupled to drive thesecond compression section.

In a further embodiment of any of the foregoing turbofan engines, the atleast two stages include a single turbine section forward of the fandrive turbine.

In a further embodiment of any of the foregoing turbofan engines, thefan section provides a fan pressure ratio between about 1.45 and about1.20.

In a further embodiment of any of the foregoing turbofan engines, thefirst compression section includes three stages and the secondcompression section includes eight stages.

Although the different examples have the specific components shown inthe illustrations, embodiments of this disclosure are not limited tothose particular combinations. It is possible to use some of thecomponents or features from one of the examples in combination withfeatures or components from another one of the examples.

These and other features disclosed herein can be best understood fromthe following specification and drawings, the following of which is abrief description.

BRIEF DESCRIPTION OF THE DRAWINGS

FIG. 1 is a schematic view of an example turbine engine according to anembodiment.

FIG. 2 is a schematic view of a compressor section of the exampleturbine engine according to an embodiment.

FIG. 3 is a schematic view of another compressor section of the exampleturbine engine according to an embodiment.

DETAILED DESCRIPTION

FIG. 1 schematically illustrates an example gas turbine engine 20 thatincludes a fan section 22, a compressor section 24, a combustor section26 and a turbine section 28. Alternative engines might include anaugmenter section (not shown) among other systems or features. The fansection 22 drives air along a bypass flow path B while the compressorsection 24 draws air in along a core flow path C where air is compressedand communicated to a combustor section 26. In the combustor section 26,air is mixed with fuel and ignited to generate a high pressure exhaustgas stream that expands through the turbine section 28 where energy isextracted and utilized to drive the fan section 22 and the compressorsection 24.

Although the disclosed non-limiting embodiment depicts a turbofan gasturbine engine, it should be understood that the concepts describedherein are not limited to use with turbofans as the teachings may beapplied to other types of turbine engines; for example a turbine engineincluding a three-spool architecture in which three spoolsconcentrically rotate about a common axis and where a low spool enablesa low pressure turbine to drive a fan via a gearbox, an intermediatespool that enables an intermediate pressure turbine to drive a firstcompressor of the compressor section, and a high spool that enables ahigh pressure turbine to drive a high pressure compressor of thecompressor section.

The example engine 20 generally includes a low speed spool 30 and a highspeed spool 32 mounted for rotation about an engine central longitudinalaxis A relative to an engine static structure 36 via several bearingsystems 38. It should be understood that various bearing systems 38 atvarious locations may alternatively or additionally be provided.

The low speed spool 30 generally includes an inner shaft 40 thatconnects the fan section 22 and a low pressure (or first) compressorsection 44 to a low pressure (or first) turbine section 46. The innershaft 40 drives the fan section 22 through a speed change device, suchas a geared architecture 48, to drive the fan section 22 at a lowerspeed than the low speed spool 30. The high-speed spool 32 includes anouter shaft 50 that interconnects a high pressure (or second) compressorsection 52 and a high pressure (or second) turbine section 54. The innershaft 40 and the outer shaft 50 are concentric and rotate via thebearing systems 38 about the engine central longitudinal axis A.

A combustor 56 is arranged between the high pressure compressor 52 andthe high pressure turbine 54. In one example, the high pressure turbine54 includes at least two stages to provide a double stage high pressureturbine 54. As used herein, a “high pressure” compressor or turbineexperiences a higher pressure than a corresponding “low pressure”compressor or turbine.

The example low pressure turbine 46 has a pressure ratio that is greaterthan about 5. The pressure ratio of the example low pressure turbine 46is measured prior to an inlet of the low pressure turbine 46 as relatedto the pressure measured at the outlet of the low pressure turbine 46prior to an exhaust nozzle. The low pressure turbine 46 is coupled tothe fan section 22 through the geared architecture 48 and therefore isalso referred to interchangeably in this disclosure as the fan driveturbine 46.

A mid-turbine frame 58 of the engine static structure 36 is arrangedgenerally between the high pressure turbine 54 and the low pressureturbine 46. The mid-turbine frame 58 further supports bearing systems 38in the turbine section 28 as well as setting airflow entering the fandrive turbine 46.

Airflow through the core airflow path C is compressed by the lowpressure compressor 44 then by the high pressure compressor 52 mixedwith fuel and ignited in the combustor 56 to produce high speed exhaustgases that are then expanded through the high pressure turbine 54 andfan drive turbine 46. The mid-turbine frame 58 includes vanes 60, whichare in the core airflow path and function as an inlet guide vane for thelow pressure turbine 46. Utilizing the vane 60 of the mid-turbine frame58 as the inlet guide vane for low pressure turbine 46 decreases thelength of the low pressure turbine 46 without increasing the axiallength of the mid-turbine frame 58. Choosing a high gearbox input tooutput ratio, reduces the number of vane rows in the fan drive turbine46 and shortens the axial length of the turbine section 28. Thus, thecompactness of the gas turbine engine 20 is increased and a higher powerdensity may be achieved.

The disclosed gas turbine engine 20 in one example is a high-bypassgeared aircraft engine. In a further example, the gas turbine engine 20includes a bypass ratio greater than about eight (8), with an exampleembodiment being greater than about twelve (12). The geared architecture48 is an epicyclical gear train, such as a planetary gear system, stargear system or other known gear system, with a gear reduction ratio ofgreater than about 2.6.

In one disclosed embodiment, the gas turbine engine 20 includes a bypassratio greater than about twelve (12:1) and a diameter of the fan blades42 is significantly larger than an outer diameter of the low pressurecompressor 44. It should be understood, however, that the aboveparameters are only exemplary of one embodiment of a gas turbine engineincluding a geared architecture and that the present disclosure isapplicable to other gas turbine engines.

A significant amount of thrust is provided by flow through the bypassflow path B due to the high bypass ratio. The fan section 22 of theengine 20 is designed for a particular flight condition—typically cruiseat about 0.8 Mach and about 35,000 feet. The flight condition of 0.8Mach and 35,000 ft., with the engine at its best fuel consumption—alsoknown as “bucket cruise Thrust Specific Fuel Consumption (‘TSFC’)”—isthe industry standard parameter of pound-mass (lbm) of fuel per hourbeing burned divided by pound-force (lbf) of thrust the engine producesat that minimum point.

“Low fan pressure ratio” is the pressure ratio across the fan bladealone, without a Fan Exit Guide Vane (“FEGV”) system. The low fanpressure ratio as disclosed herein according to one non-limitingembodiment is less than about 1.50. In another non-limiting embodiment,the low fan pressure ratio is between 1.45 and 1.20.

“Low corrected fan tip speed” is the actual fan tip speed in ft/secdivided by an industry standard temperature correction of [(Tram°R)/(518.7 OR)]^(0.5). The “Low corrected fan tip speed”, as disclosedherein according to one non-limiting embodiment, is less than about 1150ft/second.

The example gas turbine engine includes the fan section 22 thatcomprises in one non-limiting embodiment less than about 26 fan blades42. In another non-limiting embodiment, the fan section 22 includes lessthan about 20 fan blades 42. Moreover, in one disclosed embodiment thelow pressure turbine 46 includes no more than about 6 turbine rotorstages schematically indicated at 34. In another non-limiting exampleembodiment, the low pressure turbine 46 includes about 3 turbine rotorstates. A ratio between the number of fan blades 42 and the number oflow pressure turbine rotors is between about 3.3 and about 8.6. Theexample low pressure turbine 46 provides the driving power to rotate thefan section 22 and therefore the relationship between the number ofturbine rotor stages 34 in the low pressure turbine 46 and the number ofblades 42 in the fan section 22 disclose an example gas turbine engine20 with increased power transfer efficiency.

An example disclosed engine 20 provides a system-level combination ofcomponent (module) efficiencies and a system-level combination offeatures within these modules that are used to arrive at uniquely highengine efficiency (i.e. Thrust Specific Fuel Consumption) at takeoff andat bucket cruise. The disclosed combination of components providebenefit in a commercial engine with very high bypass ratio in achievingthe stated, very low, thrust specific fuel consumption (see table 1) andis especially beneficial to a single aisle aircraft where the overallpressure ratio of the compressor is less than 50.

TABLE 1 Sea level [2]Sea level takeoff, takeoff, 86 deg F., 86 deg F.,0.0 Mn: 0.0 Mn: Test Stand Test Stand Operation: Operation: Bucket nopower no power Cruise, extraction, no extraction, no 0.8 Mn,Environmental Environmental 35,000 ft, Control Control Standard Systembleed System bleed Day Thrust Specific Fuel 0.2751 0.53717 Consumption[1] Speed change At least 2.6 2.6 (Input/output) Component efficiencyFan OD at least 0.90 0.9344 0.96501 Speed Change at least 0.985 0.99490.99374 Mechanism First at least 0.84 0.8695 0.86622 Compressor or LPCSecond at least 0.82 0.8495 0.8356 Compressor or HPC Turbine at least0.85 0.87544 0.8938 Section(s) for the single excluding the HPT orcombined fan drive efficiency if turbine two or more turbines are usedFan Drive turbine at least 0.89 0.9251 0.9266

The combination of module efficiency includes among other possiblethings, the fan section 22 with the fan blades 42 supported on a fan hub64. Each of the fan blades 42 includes a leading edge 62 that extends aradial distance 66 from the engine axis A. The fan hub 64 extends aradial distance 68 from the engine axis A. A low hub-tip ratio of fanhub radial radius 68 to the radius at the leading edge 62 of the fanblade 42 is less than 0.34 and greater than 0.020. The disclosed rangeof ratios is desirable in that the lower this value is, the smaller theouter fan section and inlet section has to be to accommodate a givenamount of air, and maintaining this dimension within the desired ratiorange enables a reduction in engine weight relative to an engine with ahigher hub to tip ratio. In one example embodiment, the fan section 22further provides a low fan pressure ratio that is between about 1.45 andabout 1.20, and a bypass ratio greater than about 8.0.

The disclosed engine 20 includes the geared architecture 48 with a gearratio greater than about 2.6 to 1. In this example, the speed changesystem is the geared architecture, which is an epicyclical gearbox andwhich includes planet gears or star gears interspersed by baffles forgathering and directing lubricant during operation.

The example turbine section 28 has at least two turbine stages forwardof the first turbine stage 94 included in the fan drive turbine 46. Inthis example, the high pressure or second turbine includes two turbinestages 96 forward of the fan drive turbine 46. In this example, the twoturbine stages 96 are part of a single high pressure turbine 54 with atleast two turbine rotors 96, however, it is within the contemplation ofthis disclosure that the at least two turbine rotors forward of the fandrive turbine 46 could be part of multiple turbines that rotateindependent of each other, for example, two separate turbine sectionswith at least one turbine rotor each.

Referring to FIG. 2, with continued reference to FIG. 1, the firstcompression section 44, which in one disclosed example is a low pressurecompressor (LPC) 44, includes three stages. The example LPC 44 includesfirst compressor blade 70 with a leading edge 72 and a last compressorblade 74 with trailing edge 76. A tip of the leading edge 72 of thefirst blade 70 extends a radial distance 78 from the engine axis A. Atip of the trailing edge 76 of the last blade 74 extends a radialdistance 80 from the engine axis A. The first compression section 44 isconfigured such that a ratio between the radial distance 80 at thetrailing edge 76 is greater than 67% of the radial distance 78 of theleading edge 72 of the first blade 70. The example configurationprovided by the disclosed ratio enables improved airflow through thefirst compressor section 44 that provides improved efficiency. Thedisclosed relationship between the leading edge 72 and the trailing edge76 enables a beneficial modest slope to the engine casing structuresspanning the compressor section 24. The modest slope provides forminimal effects to tip clearances of the compressor blades due to axialshifting of the compressor rotor due to overall aerodynamic loading.

Referring to FIG. 3 with continued reference to FIG. 1, the secondcompression section 52, which in one disclosed example is a highpressure compressor (HPC), includes at least eight stages. The exampleHPC 52 includes a first blade 82 with a leading edge 84 that extends aradial distance 86 from the engine axis A to a tip. The secondcompressor section 52 also includes a last blade 88 having a trailingedge 90 that extends a radial distance 92 from the engine axis A to thetip. A ratio between the leading edge 84 and the trailing edge 90defines the configuration of the compressor 52 that provides theimproved efficiency. In one disclosed example, the radial distance 92 ofthe trailing edge 90 of the last blade 88 is greater than about 57% ofthe radial distance 86 of the leading edge 84 of the first blade 82 ofthe second compressor section 52.

A geared turbofan arrangement for short range aircraft can uniquelyexploit the particular aspects of an aircraft duty cycle that ischaracterized by an unusually low proportion of time in cruise operationversus the total time spent at takeoff and climb power (for arepresentative time span such as between engine overhauls).

A definition of a short range aircraft is one with a total flight lengthless than about 300 nautical miles.

TABLE 2 Engine #1 #2 Max takeoff 53,060 kg (117,000 lb) 58,967 kg(130,000 lb) weight Max landing 49,895 kg (110,000 lb) 55,111 kg(121,500 lb) weight Maximum 3,629 kg (8,000 lb) 4,853 kg (10,700 lb)cargo payload Maximum payload 13,676 kg (30,150 lb) 16,284 kg (35,900lb) (total) Max range 2,778 km (1,500 nmi) 2,778 km (1,500 nmi) Take off1,219 m (3,999 ft) 1,524 m (5,000 ft) run at MTOW Landing field 1,341 m(4,400 ft) 1,448 m (4,751 ft) length at MLW

As is shown in Table 2, a short range aircraft for purposes of thisdisclosure is defined as including a single aisle configuration with 2,3 seating or 3, 3 seating. Conventionally, a short range aircraft have acapacity of about 200 passengers or less. Moreover, an example shortrange aircraft will have a maximum range of only about 1500 nauticalmiles.

Because of the extremely high utilization in terms of cumulative hoursat relatively high power during take-off conditions, the disclosedgeared turbofan engine 20 arrangement is configured differently toachieve a beneficial balance of fuel burn and maintenance costs. Thehigh power utilization is a result of frequent operation at high powerconditions that generate high turbine temperatures, elevated turbinecooling air temperatures and elevated temperatures at the rear stage ofthe compressor. The result of such operation is that LPC pressures rise,temperature rise and efficiency may be lower than for a long rangeaircraft. In a long range aircraft that operates for longer periods anda greater portion of the cumulative operating hours, maximizing LPCefficiency is desired provides a significant benefit, and is a keydifference when compared to short range aircraft. Pressure andtemperature rise can be increased due to the less frequent use oftakeoff power between overhaul periods which could be around 4000 hoursfor both the short range and long range commercial airliner.

Although an example embodiment has been disclosed, a worker of ordinaryskill in this art would recognize that certain modifications would comewithin the scope of this disclosure. For that reason, the followingclaims should be studied to determine the scope and content of thisdisclosure.

What is claimed is:
 1. A turbofan engine comprising; a fan sectionincluding a fan blade having a leading edge and a hub to tip ratio ofless than 0.34 and greater than 0.020 measured at the leading edge; afirst compression section; a second compression section including a lastblade trailing edge radial tip length that is greater than 57% of theradial tip length of the leading edge of the first stage of the secondcompression section; a turbine section including a fan drive turbinesection and at least two turbine stages forward of a first turbine bladeof the fan drive turbine section; and a mid-turbine frame disposedbetween the fan drive turbine and the at least two turbine stages forguiding core flow into the fan drive turbine, the mid-turbine framesupporting a bearing system for one of the fan drive turbine and the atleast two turbine stages.
 2. The turbofan engine as recited in claim 1,wherein the first compression section includes a last blade trailingedge radial tip length that is greater than 67% of a radial tip lengthof a leading edge of a first stage of the first compression section. 3.The turbofan engine as recited in claim 2, including a speed changesystem coupled between the fan drive turbine and the fan section, thespeed change system including a gear ratio greater than 2.6 to
 1. 4. Theturbofan engine as recited in claim 3, wherein the fan section providesa low fan pressure ratio less than 1.6.
 5. The turbofan engine asrecited in claim 3, wherein the fan section provides a low fan pressureratio between 1.45 and 1.20.
 6. The turbofan engine as recited in claim3, wherein the fan section provides a bypass ratio greater than
 8. 7.The turbofan engine as recited in claim 3, wherein the fan sectionprovides a bypass ratio greater than
 12. 8. The turbofan engine asrecited in claim 5, wherein the fan drive turbine includes at leastthree stages.
 9. The turbofan engine as recited in claim 1, wherein thefan drive turbine is coupled to drive the first compression section. 10.The turbofan engine as recited in claim 9, wherein the at least twostages comprise a two stage turbine forward of the fan drive turbine.11. The turbofan engine as recited in claim 9, wherein the fan driveturbine comprises a turbine with from three to six stages.
 12. Aturbofan engine comprising: a fan section providing a bypass ratiogreater than 8, the fan section includes a plurality of fan bladessupported on a hub; a speed change system coupled to the hub with a gearratio greater than about 2.6 to 1; a first compression section includinga last blade trailing edge tip length that is greater than 67% of theradial length of a first stage leading edge of the first compressionsection, wherein the first compression section operates at an efficiencyof at least 0.84; a turbine section including a fan drive turbinesection coupled to the speed change system and at least two turbinestages forward of a first turbine blade of the fan drive turbinesection; and a mid-turbine frame disposed between the fan drive turbineand the at least two turbine stages for guiding core flow into the fandrive turbine, the mid-turbine frame supporting a bearing system for oneof the fan drive turbine and the at least two turbine stages.
 13. Theturbofan engine as recited in claim 12, wherein at least one of theplurality of fan blades includes a leading edge and a hub to tip ratiois less than about 0.34 and greater than about 0.020 measured at theleading edge.
 14. A turbofan engine as recited in claim 13, including aneight stage second compression section including a last blade with atrailing edge tip that includes a radial length that is greater than 57%of a radial length of the first stage leading edge of the secondcompression section.
 15. The turbofan engine as recited in claim 14,wherein the second compression section operates at an efficiency of atleast 0.82.
 16. The turbofan engine as recited in claim 14, wherein theat least two stages comprise a two-stage second turbine section forwardof the fan drive turbine.
 17. The turbofan engine as recited in claim14, wherein the fan drive turbine operates at an efficiency of at least0.89 and the at least two turbine stages forward of the fan driveturbine section comprises a second turbine section operating at anefficiency of at least 0.85.
 18. The turbofan engine as recited in claim17, wherein the fan section includes a fan case circumscribing theplurality of fan blades and defining a bypass flow path.
 19. Theturbofan engine as recited in claim 18, wherein the fan drive turbineincludes a plurality of rotor stages and a ratio between the number ofthe plurality of fan blades and the number of the plurality of rotorstages is between 3.3 and 8.6.
 20. A turbofan engine comprising: a fansection providing a bypass ratio greater than 8, the fan sectionincludes a plurality of fan blades supported on a hub, wherein at leastone of the plurality of fan blades includes a leading edge and a hub totip ratio is less than about 0.34 and greater than about 0.020 measuredat the leading edge; a three-stage first compression section disposedaft of the speed change system, the three-stage first compressionsection including a first rotating compressor blade with a leading edgeand a last rotating compressor blade with a trailing edge, wherein aradial distance from an engine longitudinal axis to a tip of thetrailing edge of the last rotating compressor blade is greater than 67%of a radial distance from the engine longitudinal axis to a tip of thefirst rotating compressor blade at the leading edge; an eight stagesecond compression section including a last rotating compressor bladewith a trailing edge tip that includes a radial length that is greaterthan 57% of a radial length of the first rotating compressor blade atthe leading edge of the eight-stage first compression section; a turbinesection including a fan drive turbine section and a two stage turbinesection forward of a first turbine blade of the fan drive turbinesection; and a geared architecture coupled between the fan drive turbineand the hub of the fan section, the geared architecture including with agear ratio greater than about 2.6 to
 1. 21. The turbofan engine asrecited in claim 20, wherein the fan drive turbine operates at anefficiency of at least 0.89.
 22. The turbofan engine as recited in claim20, including a mid-turbine frame disposed between the fan drive turbineand the two stage turbine section for guiding core flow into the fandrive turbine, the mid-turbine frame supporting a bearing system for oneof the fan drive turbine and the two stage turbine section.
 23. Theturbofan engine as recited in claim 22, wherein the mid-turbine frameincludes a vane guiding a core flow into the fan drive turbine section.24. The turbofan engine as recited in claim 23, wherein the firstcompression section includes an efficiency of at least 0.84 and thesecond compression section includes an efficiency of at least 0.82. 25.The turbofan engine as recited in claim 24, wherein the fan driveturbine operates at an efficiency of at least 0.89.
 26. The turbofanengine as recited in claim 25, wherein the two stage turbine operates atan efficiency of at least 0.85.
 27. The turbofan engine as recited inclaim 25, wherein the speed change system comprises an epicyclic gearsystem that operates at an efficiency of at least 0.985.
 28. Theturbofan engine as recited in claim 27, wherein the fan includes lessthan 26 fan blades.
 29. The turbofan engine as recited in claim 28,wherein a ratio between the number of the plurality of fan blades andthe number of the plurality of stages of the fan drive turbine isbetween 3.3 and 8.6.
 30. The turbofan engine as recited in claim 29,wherein the fan section provides a low fan pressure ratio between 1.45and 1.20.